Aviation hybrid engine with electric drive unit thermal management system

Abstract

The invention provides an aviation hybrid power engine with an electric drive unit thermal management system, which comprises an aircraft fuel tank, an air fuel radiator and the electric drive unit thermal management system, and the electric drive unit thermal management system is independent of a fuel flow path of the aviation hybrid power engine. The electric drive unit heat management system is connected between the air fuel oil radiator and the aircraft fuel oil tank in series, and cooling fuel oil is continuously provided for the electric drive unit through the electric drive unit heat management system. According to the invention, the cooling medium with appropriate flow pressure and low temperature can be provided for the propulsion motor stator and the propulsion motor controller, so that the temperature and pressure requirements of the propulsion motor stator and the propulsion motor controller can be met, and the heat generation amount of the propulsion motor stator and the propulsion motor controller can be absorbed; normal operation of the propulsion motor and the motor controller is guaranteed, and the safety of the aviation hybrid power engine is improved.

Description

Technical Field

[0001] This invention relates to a thermal management system for aviation hybrid engines, and particularly to an aviation hybrid engine with an electric drive unit thermal management system. Background Technology

[0002] Guided by goals such as carbon neutrality and green aviation, the aviation industry faces increasingly stringent emission and economic requirements. Therefore, it has begun to focus on the demand for hybrid power technology with new energy configurations, and is developing aviation hybrid power engines.

[0003] The electric drive unit in a hybrid power system includes a propulsion motor and a propulsion motor controller. Due to the high power of the electric drive unit, the propulsion motor and controller generate significant heat. Furthermore, the propulsion motor is generally allowed to operate at a temperature not exceeding 200°C, and the propulsion motor controller is generally not allowed to operate at a temperature exceeding 100°C.

[0004] Therefore, a cooling medium is needed to dissipate heat from the propulsion motor and its controller, and a thermal management system architecture needs to be designed. Since most of the heat in the propulsion motor is generated by the stator, the commonly used thermal management system for propulsion motors involves immersing the stator in an insulating cooling medium. The propulsion motor controller is designed with cooling medium channels, through which the cooling medium carries away the heat from the controller.

[0005] Currently, the main problems with thermal management of the electric drive unit in aircraft hybrid power engines are as follows:

[0006] I. The propulsion motors of hybrid-electric aircraft engines have high power outputs, reaching megawatt levels or higher. These high-power propulsion motors generate significant heat during operation, causing a rapid rise in temperature. Excessive heat can lead to coking of the cooling medium and demagnetization of the permanent magnets, affecting the motor's safety and lifespan. The propulsion motor controller also generates considerable heat during operation, and some of its components are temperature-sensitive, thus requiring high-temperature cooling media.

[0007] II. Lubricating oil and fuel are commonly used cooling media in aero engines. In the engine's fuel flow path and lubricating oil flow path, lubricating oil is used to lubricate and cool bearings, seals, splines, and gears through the oil supply subsystem, oil return subsystem, and ventilation subsystem, and to ensure the circulation of lubricating oil within the engine.

[0008] After fuel is supplied to the engine from the aircraft, it is first pressurized by the fuel pump, and then used by the fuel metering device and fuel distribution device to achieve functions such as fuel supply to the combustion chamber and actuation of the servo device. It also exchanges heat with lubricating oil through the fuel-oil heat exchanger.

[0009] However, if the fuel and lubricating oil in the prime mover's fuel flow path and lubricating oil flow path are used as cooling media in the thermal management system of the electric drive unit, the following problems will occur:

[0010] Using lubricating oil as the cooling medium requires an additional oil tank, increasing the overall weight of the engine and reducing fuel economy. Moreover, under certain high-temperature operating conditions, the lubricating oil supply temperature is high, and the heat transfer capacity of the lubricating oil cooler is limited, making it difficult to cool the lubricating oil to the limiting temperature of the cooling medium inlet of the propulsion motor controller, thus failing to meet the heat dissipation requirements of the propulsion motor controller.

[0011] Using prime mover fuel as the cooling medium results in a high fuel pressure after the fuel pump is boosted, which is difficult to meet the pressure resistance requirements of the propulsion motor and propulsion motor controller. At the same time, it increases the risk of cooling medium leakage in the electric drive unit, affecting the safety of the electric drive unit.

[0012] Third, in aero-engine hybrid power systems, the operating conditions of the prime mover and the propulsion motor are not perfectly matched. If the lubricating oil system of the prime mover is used to cool the electric drive unit, then under certain operating conditions, the flow rate in the lubricating oil path and fuel path of the prime mover is relatively low, while the power of the propulsion motor and propulsion motor controller is relatively high. As a result, the fuel or lubricating oil heat sink cannot completely remove the heat generated by the propulsion motor and propulsion motor controller.

[0013] When the propulsion motor works independently, the prime mover stops, the oil pump driven by the prime mover stops working, and the fuel and lubricating oil in the prime mover stop flowing. However, the flow path of the thermal management system of the propulsion motor and the propulsion motor controller still needs to work normally.

[0014] In view of this, the inventors of this application have designed an aviation hybrid power engine with an electric drive unit thermal management system in order to overcome the above-mentioned technical problems. Summary of the Invention

[0015] The technical problem to be solved by the present invention is to overcome the defects of the prior art, which uses fuel lubricating oil as a cooling medium, and the high fuel pressure after the fuel pump is boosted is difficult to meet the pressure resistance requirements of the propulsion motor and propulsion motor controller, which increases the risk of cooling medium leakage in the electric drive unit and affects the safety of the electric drive unit. The present invention provides an aviation hybrid power engine with an electric drive unit thermal management system.

[0016] The present invention solves the above-mentioned technical problems through the following technical solution:

[0017] An aviation hybrid power engine with an electric drive unit thermal management system is characterized in that the aviation hybrid power engine includes an aircraft fuel tank, an air-fuel radiator, and an electric drive unit thermal management system. The electric drive unit thermal management system is independent of the fuel flow path of the aviation hybrid power engine. The electric drive unit thermal management system is connected in series between the air-fuel radiator and the aircraft fuel tank, and continuously provides cooling fuel to the electric drive unit through the electric drive unit thermal management system.

[0018] According to one embodiment of the present invention, the electric drive unit thermal management system includes an electric oil pump, a propulsion motor controller cooling medium channel, and a propulsion motor stator cooling medium channel connected in series. The electric oil pump is connected to the aircraft fuel tank, and the propulsion motor stator cooling medium channel is connected to the air-fuel radiator.

[0019] According to one embodiment of the present invention, a fuel filter is connected between the air fuel radiator and the aircraft fuel tank.

[0020] According to one embodiment of the present invention, the fuel outlet of the aircraft fuel tank is connected to the fuel inlet of the electric fuel pump via a fluid line.

[0021] According to one embodiment of the present invention, the fuel outlet of the electric oil pump is connected to the fuel inlet of the cooling medium channel of the propulsion motor controller via a fluid pipeline.

[0022] According to one embodiment of the present invention, the fuel outlet of the propulsion motor controller cooling medium channel is connected to the fuel inlet of the propulsion motor stator cooling medium channel via a fluid pipeline.

[0023] According to one embodiment of the present invention, the fuel outlet of the stator cooling medium flow channel of the propulsion motor is connected to the fuel inlet of the air-fuel radiator via a fluid pipeline.

[0024] According to one embodiment of the present invention, the fuel outlet of the air-fuel radiator is connected to the fuel inlet of the fuel filter via a fluid pipeline, and the fuel outlet of the fuel filter is connected to the fuel inlet of the aircraft fuel tank, so that the cooled fuel flows back to the aircraft fuel tank.

[0025] According to one embodiment of the present invention, the aero-hybrid engine further includes an open rotor fan, a propulsion motor, a prime mover compressor, and an engine turbine, wherein the open rotor fan, the propulsion motor, the air intake portion of the prime mover compressor, and the low-pressure turbine portion of the engine turbine are coaxially connected.

[0026] According to one embodiment of the present invention, the aero-hybrid engine further includes a prime mover compressor, a prime mover turbine, a prime mover housing, and a propulsion motor controller, wherein the high-pressure compressor portion of the prime mover compressor and the high-pressure turbine portion of the prime mover turbine are coaxially connected.

[0027] The air-fuel radiator and the propulsion motor controller are located between the prime mover housing and the engine turbine.

[0028] The positive and progressive effects of this invention are as follows:

[0029] The present invention relates to an aero-hybrid engine with an electric drive unit thermal management system, which can ensure that a cooling medium with suitable flow rate, pressure and low temperature is provided to the propulsion motor stator and propulsion motor controller, thereby meeting the temperature and pressure requirements of the propulsion motor stator and propulsion motor controller, and absorbing the heat generated by the propulsion motor stator and propulsion motor controller, ensuring the normal operation of the propulsion motor and motor controller, and improving the safety of the aero-hybrid engine.

[0030] Meanwhile, the heated fuel, when returned to the aircraft's fuel tank, can increase the temperature of the fuel inside the tank and reduce the risk of fuel freezing. Attached Figure Description

[0031] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, in which the same reference numerals always denote the same features, wherein:

[0032] Figure 1 This is a schematic diagram of the architecture of the electric drive unit thermal management system in the aero-hybrid engine with electric drive unit thermal management system of the present invention.

[0033] Figure 2 This is a schematic diagram of the structure of the aviation hybrid power engine with an electric drive unit thermal management system according to the present invention. Detailed Implementation

[0034] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0035] Embodiments of the invention will now be described in detail with reference to the accompanying drawings. Preferred embodiments of the invention will now be described in detail, examples of which are shown in the drawings. Wherever possible, the same reference numerals will be used in all the drawings to denote the same or similar parts.

[0036] Furthermore, although the terminology used in this invention is selected from commonly known and used terms, some terms mentioned in this specification may have been selected by the applicant in his or her judgment, and their detailed meanings are explained in the relevant sections of the description herein.

[0037] Furthermore, the invention should be understood not only through the actual terminology used, but also through the meaning implied by each term.

[0038] Figure 1 This is a schematic diagram of the architecture of the electric drive unit thermal management system in the aero-hybrid engine with electric drive unit thermal management system of the present invention.

[0039] like Figure 1 As shown, this invention discloses an aero-hybrid engine with an electric drive unit thermal management system, comprising an aircraft fuel tank 10, an air-fuel radiator 20, and an electric drive unit thermal management system 30. The electric drive unit thermal management system 30 is independent of the fuel flow path of the aero-hybrid engine and is connected in series between the air-fuel radiator 20 and the aircraft fuel tank 10, continuously providing cooling fuel to the electric drive unit.

[0040] Preferably, the electric drive unit thermal management system 30 includes an electric oil pump 31, a propulsion motor controller cooling medium channel 32, and a propulsion motor stator cooling medium channel 33 connected in series. The electric oil pump 31 is connected to the aircraft fuel tank 10, and the propulsion motor stator cooling medium channel 33 is connected to the air-fuel radiator 20. A fuel filter 34 is connected between the air-fuel radiator 20 and the aircraft fuel tank 10.

[0041] The fuel outlet of the aircraft fuel tank 10 is connected to the fuel inlet of the electric fuel pump 31 via a fluid line. The electric fuel pump 31 controls the fuel flow rate according to the power of the propulsion motor and the propulsion motor controller. The fuel outlet of the electric fuel pump 31 is connected to the fuel inlet of the propulsion motor controller cooling medium channel 32 via a fluid line. The fuel outlet of the propulsion motor controller cooling medium channel 32 is connected to the fuel inlet of the propulsion motor stator cooling medium channel 33 via a fluid line.

[0042] Fuel enters the cooling medium channel 32 of the propulsion motor controller and the cooling medium channel 33 of the propulsion motor stator successively, absorbing the heat generated by the motor controller and the propulsion motor.

[0043] The fuel outlet of the stator cooling medium flow channel 33 of the propulsion motor is connected to the fuel inlet of the air-fuel radiator 20 via a fluid pipeline. The fuel outlet of the air-fuel radiator 20 is connected to the fuel inlet of the fuel filter 34 via a fluid pipeline, and the fuel outlet of the fuel filter 34 is connected to the fuel inlet of the aircraft fuel tank 10, so that the cooled fuel flows back to the aircraft fuel tank 10.

[0044] To ensure that the fuel temperature does not exceed the temperature limit for fuel return from the aircraft fuel tank, the fuel needs to be cooled by air through an air-fuel radiator and then filtered through a fuel filter before returning to the aircraft fuel tank.

[0045] Figure 2 This is a schematic diagram of the structure of the aviation hybrid power engine with an electric drive unit thermal management system according to the present invention.

[0046] like Figure 2 As shown, the aero-hybrid engine includes an open rotor fan 100, a prime mover compressor 110, a prime mover housing 120, a prime mover combustion chamber 130, an air-fuel radiator 140, a prime mover turbine 150, a prime mover exhaust nozzle 160, a propulsion motor 170, and a propulsion motor controller 180. The air intake sections of the open rotor fan 100, propulsion motor 170, and prime mover compressor 110 are coaxially connected to the low-pressure turbine section of the engine turbine 150. The high-pressure compressor section of the prime mover compressor 110 is coaxially connected to the high-pressure turbine section of the prime mover turbine 150. The air-fuel radiator 140 and propulsion motor controller 180 are located between the prime mover housing 120 and the engine turbine 150.

[0047] During the process of fuel consumption by the prime mover, the aircraft fuel tank needs to continuously supply fuel to the engine. This fuel has a relatively low temperature, which meets the cooling medium temperature requirements of the electric drive unit. At the same time, the pressure can meet the pressure resistance requirements and flow resistance of the propulsion motor and propulsion motor controller.

[0048] This invention relates to an aero-hybrid engine with an electric drive unit thermal management system. An electric fuel pump independently supplies fuel to the electric drive unit; specifically, an additional portion of fuel is supplied from the aircraft fuel tank into the cooling medium channels of the propulsion motor controller and stator to absorb the generated heat. Subsequently, the fuel temperature is reduced to the permissible return temperature of the aircraft fuel tank via an air-fuel radiator, and finally returned to the aircraft fuel tank.

[0049] The present invention relates to an aero-hybrid engine with an electric drive unit thermal management system, which has the following advantages:

[0050] I. This invention considers the integrated design of the fuel thermal management system for aircraft and aviation hybrid power engines, realizing integrated thermal management of the aircraft and engine, improving the temperature of the fuel in the fuel tank and reducing the risk of fuel icing.

[0051] Second, the aviation hybrid power engine of the present invention draws fuel-cooled electric drive units from the aircraft fuel tank, which not only meets the temperature and pressure requirements of the cooling medium of the electric drive units, but also eliminates the need for an additional fuel tank, thus reducing the weight of the engine.

[0052] Third, the electric drive unit thermal management system adopted in this invention has a relatively simple structure because the power of the propulsion motor and the propulsion motor controller are matched, and the cooling medium flow channel of the propulsion motor controller and the cooling medium flow channel of the propulsion motor stator are connected in series. It is also easier to control the fuel flow using an electric oil pump, which reduces the system complexity and improves the system reliability.

[0053] Fourth, in aero-electric hybrid engines, an open rotor configuration is used, eliminating the fan nacelle. The radiator must be installed inside the prime mover compartment, resulting in reduced airflow to the radiator. Under certain operating conditions, the radiator may struggle to completely dissipate the heat absorbed by the electric drive unit from the fuel.

[0054] The electric drive unit thermal management system used in this invention can increase fuel flow and increase fuel heat sink by adjusting the electric fuel pump, ensuring that the fuel temperature does not exceed the allowable return temperature of the aircraft fuel tank after absorbing the heat generated by the electric drive unit.

[0055] Fifth, in the case where the working states of the prime mover and the propulsion motor are not perfectly matched, the thermal management system scheme for the electric drive unit proposed in this invention continuously provides cooling fuel to the electric drive unit through an electric oil pump, thus solving the fuel supply problem for the electric drive unit when it operates alone.

[0056] 6. In response to the risk of coking of the fuel cooling medium in the electric drive unit due to heat refluxing after the engine is stopped, the electric drive unit thermal management system can continuously supply cooling fuel to the electric drive unit through an electric oil pump after the engine is stopped until the electric drive unit cools down to a safe temperature, thus solving the problem of fuel coking caused by heat refluxing.

[0057] In summary, the aviation hybrid engine with an electric drive unit thermal management system of the present invention can ensure that a cooling medium with suitable flow rate, pressure and low temperature is provided to the propulsion motor stator and propulsion motor controller, thereby meeting the temperature and pressure requirements of the propulsion motor stator and propulsion motor controller, absorbing the heat generated by the propulsion motor stator and propulsion motor controller, ensuring the normal operation of the propulsion motor and motor controller, and improving the safety of the aviation hybrid engine.

[0058] Meanwhile, the heated fuel, when returned to the aircraft's fuel tank, can increase the temperature of the fuel inside the tank and reduce the risk of fuel freezing.

[0059] For those skilled in the art, the above disclosure is merely illustrative and does not constitute a limitation of this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to this application. Such modifications, improvements, and corrections are suggested in this application and therefore remain within the spirit and scope of the exemplary embodiments of this application.

[0060] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0061] Similarly, it should be noted that, in order to simplify the description of the embodiments disclosed in this application and thus aid in the understanding of one or more embodiments of the invention, the foregoing description of the embodiments of this application sometimes combines multiple features into a single embodiment, drawing, or description thereof. However, this disclosure method does not imply that the subject matter of this application requires more features than those mentioned in the claims. In fact, the embodiments have fewer features than all the features of the single embodiments disclosed above.

[0062] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. An aero-electric hybrid power engine with an electric drive unit thermal management system, characterized in that, The aero-hybrid engine includes an aircraft fuel tank, an air-fuel radiator, and an electric drive unit thermal management system. The electric drive unit thermal management system is independent of the fuel flow path of the aero-hybrid engine. The electric drive unit thermal management system is connected in series between the air-fuel radiator and the aircraft fuel tank, and continuously provides cooling fuel to the electric drive unit.

2. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 1, characterized in that, The electric drive unit thermal management system includes an electric oil pump, a propulsion motor controller cooling medium channel, and a propulsion motor stator cooling medium channel connected in series. The electric oil pump is connected to the aircraft fuel tank, and the propulsion motor stator cooling medium channel is connected to the air-fuel radiator.

3. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 2, characterized in that, A fuel filter is connected between the air fuel radiator and the aircraft fuel tank.

4. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 2, characterized in that, The fuel outlet of the aircraft fuel tank is connected to the fuel inlet of the electric fuel pump via a fluid line.

5. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 4, characterized in that, The fuel outlet of the electric oil pump is connected to the fuel inlet of the cooling medium channel of the propulsion motor controller via a fluid pipeline.

6. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 5, characterized in that, The fuel outlet of the propulsion motor controller cooling medium channel is connected to the fuel inlet of the propulsion motor stator cooling medium channel via a fluid pipeline.

7. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 6, characterized in that, The fuel outlet of the stator cooling medium flow channel of the propulsion motor is connected to the fuel inlet of the air-fuel radiator via a fluid pipeline.

8. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 7, characterized in that, The fuel outlet of the air-fuel radiator is connected to the fuel inlet of the fuel filter via a fluid pipeline, and the fuel outlet of the fuel filter is connected to the fuel inlet of the aircraft fuel tank, so that the cooled fuel flows back to the aircraft fuel tank.

9. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 1, characterized in that, The aero-hybrid engine also includes an open rotor fan, a propulsion motor, a prime mover compressor, and an engine turbine. The open rotor fan, the propulsion motor, the air intake section of the prime mover compressor, and the low-pressure turbine section of the engine turbine are coaxially connected.

10. The aero-hybrid engine with an electric drive unit thermal management system as described in claim 9, characterized in that, The aero-hybrid engine also includes a prime mover compressor, a prime mover turbine, a prime mover housing, and a propulsion motor controller. The high-pressure compressor section of the prime mover compressor and the high-pressure turbine section of the prime mover turbine are coaxially connected. The air-fuel radiator and the propulsion motor controller are located between the prime mover housing and the engine turbine.

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